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At least 307 records · Page 17

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NASA NTRS2008

100-LBF LO2/LCH4 - Reaction Control Engine Technology Development for Future Space Vehicles

The National Aeronautics and Space Administration (NASA) has identified liquid oxygen (LO2)/liquid methane (LCH4) propulsion systems as promising options for some future space vehicles. NASA issued a contract to Aerojet to develop a 100-lbf (445 N) LO2/LCH4 Reaction Control Engine (RCE) aimed at reducing the risk of utilizing a cryogenic reaction control system (RCS) on a space vehicle. Aerojet utilized innovative design solutions to develop an RCE that can ignite reliably over a broad range of inlet temperatures, perform short minimum impulse bits (MIB) at small electrical pulse widths (EPW), and produce excellent specific impulse (Isp) across a range of engine mixture ratios (MR). These design innovations also provide a start transient with a benign MR, ensuring good thrust chamber compatibility and long life. In addition, this RCE can successfully operate at MRs associated with main engines, enabling the RCE to provide emergency backup propulsion to minimize vehicle propellant load and overall system mass.

Robinson, Philip J.↗

NASA NTRS2008

100-Lb(f) LO2/LCH4 Reaction Control Engine Technology Development for Future Space Vehicles

Robinson, Philip J.↗

NASA NTRS2012

Simulation of Oxygen Disintegration and Mixing With Hydrogen or Helium at Supercritical Pressure

The simulation of high-pressure turbulent flows, where the pressure, p, is larger than the critical value, p(sub c), for the species under consideration, is relevant to a wide array of propulsion systems, e.g. gas turbine, diesel, and liquid rocket engines. Most turbulence models, however, have been developed for atmospheric-p turbulent flows. The difference between atmospheric-p and supercritical-p turbulence is that, in the former situation, the coupling between dynamics and thermodynamics is moderate to negligible, but for the latter it is very significant, and can dominate the flow characteristics. The reason for this stems from the mathematical form of the equation of state (EOS), which is the perfect-gas EOS in the former case, and the real-gas EOS in the latter case. For flows at supercritical pressure, p, the large eddy simulation (LES) equations consist of the differential conservation equations coupled with a real-gas EOS. The equations use transport properties that depend on the thermodynamic variables. Compared to previous LES models, the differential equations contain not only the subgrid scale (SGS) fluxes, but also new SGS terms, each denoted as a correction. These additional terms, typically assumed null for atmospheric pressure flows, stem from filtering the differential governing equations, and represent differences between a filtered term and the same term computed as a function of the filtered flow field. In particular, the energy equation contains a heat-flux correction (q-correction) that is the difference between the filtered divergence of the heat flux and the divergence of the heat flux computed as a function of the filtered flow field. In a previous study, there was only partial success in modeling the q-correction term, but in this innovation, success has been achieved by using a different modeling approach. This analysis, based on a temporal mixing layer Direct Numerical Simulation database, shows that the focus in modeling the q-correction should be on reconstructing the primitive variable gradients rather than their coefficients, and proposes the approximate deconvolution model (ADM) as an effective means of flow field reconstruction for LES heat flux calculation. Further, results for a study conducted for temporal mixing layers initially containing oxygen in the lower stream, and hydrogen or helium in the upper stream, show that, for any LES, including SGS-flux models (constant-coefficient Gradient or Scale-Similarity models, dynamic-coefficient Smagorinsky/Yoshizawa or mixed Smagorinsky/Yoshizawa/Gradient models), the inclusion of the q-correction in the LES leads to the theoretical maximum reduction of the SGS heat-flux difference. The remaining error in modeling this new subgrid term is thus irreducible.

Bellan, Josette↗

NASA NTRS2017

Reduced Order Modeling of SLS Liquid Hydrogen Pre-Valve Flow Guide to Enable Rapid Transient Analysis

NASA is developing a new launch vehicle, called the Space Launch System (SLS), which is intended on taking humans out of low earth orbit to destinations including the moon, asteroids, and Mars. The propulsion system for the core stage of this vehicle includes four RS-25 Liquid Hydrogen/Oxygen rocket engines. These engines are upgraded versions of the Space Shuttle Main Engines (SSME); the upgrades include higher power levels and affordability enhancements. As with any new vehicle, the Main Propulsion System (MPS), which include the feedlines and ancillary hardware connecting the engines to the fuel and oxidizer tanks, had to be redesigned (figure 1 - export clearance in progress), as the previous MPS for the SSME's was inherently part of the Space Shuttle System, which had a completely different overall configuration.

Brown, Andrew M.↗

NASA NTRS2011

Sloshing in the Liquid Hydrogen and Liquid Oxygen Propellant Tanks After Main Engine Cut Off

NASA Marshall Space Flight Center is designing and developing the Main Propulsion System (MPS) for Ares launch vehicles. Propellant sloshing in the liquid hydrogen (LH2) and liquid oxygen (LO2) propellant tanks after Main Engine Cut Off (MECO) was modeled using the Volume of Fluid (VOF) module of the computational fluid dynamics code, CFD-ACE+. The present simulation shows that there is substantial sloshing side forces acting on the LH2 tank during the deceleration of the vehicle after MECO. The LH2 tank features a side wall drain pipe. The side loads result from the residual propellant mass motion in the LH2 tank which is initiated by the stop of flow into the drain pipe at MECO. The simulations show that radial force on the LH2 tank wall is less than 50 lbf and the radial moment calculated based up through the center of gravity of the vehicle is predicted to be as high as 300 lbf-ft. The LO2 tank features a bottom dome drain system and is equipped with sloshing baffles. The remaining LO2 in the tank slowly forms a liquid column along the centerline of tank under the zero gravity environments. The radial force on the LO2 tank wall is predicted to be less than 100 lbf. The radial moment calculated based on the center of gravity of the vehicle is predicted as high as 4500 lbf-ft just before MECO and dropped down to near zero after propellant draining stopped completely.

Kim, Sura↗

NASA NTRS1992

Advanced ramjet concepts program

Uniquely advantageous features, on both the performance and weight sides of the ledger, can be achieved through synergistic design integration of airbreathing and rocket technologies in the development of advanced orbital space transport propulsion systems of the combined cycle type. In the context of well understood advanced airbreathing and liquid rocket propulsion principles and practices, this precept of synergism is advanced mainly through six rather specific examples. These range from the detailed component level to the overall vehicle system level as follows: using jet compression; achieving a high area ratio rocket nozzle; ameliorating gas generator cycle rocket system deficiencies; using the in-duct special rocket thrust chamber assembly as the principal scramjet fuel injection operation; using the unstowed, covered fan as a duct closure for effecting high area ratio rocket mode operation; and creating a unique airbreathing rocket system via the onboard, cryogenic hydrogen induced air liquefaction process.

Leingang, J. L.↗

NASA NTRSDate not supplied

Heat Transfer Measurements of NASA Liquid Kerosene/Oxygen Rotating Detonation Rocket Engine

The RDRE has been identified as a viable high performance propulsion system with numerous advantages over the state-of-the-art (SOA) liquid rocket engine. NASA has been investigating this combustion device for applications ranging from lander, upper stage, thruster, and hypersonics. All of the activities funded to date have been focused on closing major technology gaps preventing the RDRE from being used more broadly by industry. One of those gaps include the prediction and management of heat transfer to the walls from the extreme combustion environment. This work overviews the heat transfer measurements made using new and existing hardware. A liquid oxygen / liquid RP-1 RDRE was tested utilizing a calorimeter outer body and outer body nozzle extension. An actively cooled inner body with axial running cooling channels was also used. A bimetallic GRCop-42 / Monel K500 injector was developed and demonstrated to be a viable technology for RDRE environments. Trends in bulk heat loads are discussed along with a direct comparison to constant pressure theory predictions of wall heat transfer. The experimental data obtained in this investigation showed similar heat fluxes to constant pressure theory indicating deflagration may have dominated the flow field. Only a single detonation wave was observed in all tests which imparted a significant dynamic load (vibration) on the test article making measurement of combustor performances extremely challenging. Both accelerometer and load cell data corroborate the extreme G-forces measured. The single wave mode yielded relatively low performances compared with other hot fire test data sets available. This leads to the conclusion that a single detonation wave in this geometry is not sufficient for high performance. Wave multiplicity, or rather, a specific number of waves may yield higher performances.

Thomas Teasley↗

NASA NTRS2015

Nontoxic Ionic Liquid Fuels for Exploration Applications

The toxicity of propellants used in conventional propulsion systems increases not only safety risks to personnel but also costs, due to special handling required during the entire lifetime of the propellants. Orbital Technologies Corporation (ORBITEC) has developed and tested novel nontoxic ionic liquid fuels for propulsion applications. In Phase I of the project, the company demonstrated the feasibility of several ionic liquid formulations that equaled the performance of conventional rocket propellant monomethylhydrazine (MMH) and also provided low volatility and low toxicity. In Phase II, ORBITEC refined the formulations, conducted material property tests, and investigated combustion behavior in droplet and microreactor experiments. The company also explored the effect of injector design on performance and demonstrated the fuels in a small-scale thruster. The ultimate goal is to replace propellants such as MMH with fuels that are simultaneously high-performance and nontoxic. The fuels will have uses in NASA's propulsion applications and also in a range of military and commercial functions.

Coil, Millicent↗

NASA NTRS1994

The JPL Direct Methanol Liquid-feed PEM Fuel Cell

Recently, there has been a breakthrough in fuel cell technology in the Energy Storage Systems Group at the Jet Propulsion Laboratory with the develpment of a direct methanol, liquid-feed, solid polymer electrolyte membrane (PEM) fuel cell... The methanol liquid-feed, solid polymer electrolyte (PEM) design has numerous system level advantages over the gas-feed design. These include:...

fuel cells methanol pem methanol water↗

NASA NTRS2007

Infrared Camera Characterization of Bi-Propellant Reaction Control Engines during Auxiliary Propulsion Systems Tests at NASA's White Sands Test Facility in Las Cruces, New Mexico

This paper describes the application of a FUR Systems A40M infrared (IR) digital camera for thermal monitoring of a Liquid Oxygen (LOX) and Ethanol bi-propellant Reaction Control Engine (RCE) during Auxiliary Propulsion System (APS) testing at the National Aeronautics & Space Administration's (NASA) White Sands Test Facility (WSTF) near Las Cruces, New Mexico. Typically, NASA has relied mostly on the use of ThermoCouples (TC) for this type of thermal monitoring due to the variability of constraints required to accurately map rapidly changing temperatures from ambient to glowing hot chamber material. Obtaining accurate real-time temperatures in the JR spectrum is made even more elusive by the changing emissivity of the chamber material as it begins to glow. The parameters evaluated prior to APS testing included: (1) remote operation of the A40M camera using fiber optic Firewire signal sender and receiver units; (2) operation of the camera inside a Pelco explosion proof enclosure with a germanium window; (3) remote analog signal display for real-time monitoring; (4) remote digital data acquisition of the A40M's sensor information using FUR's ThermaCAM Researcher Pro 2.8 software; and (5) overall reliability of the system. An initial characterization report was prepared after the A40M characterization tests at Marshall Space Flight Center (MSFC) to document controlled heat source comparisons to calibrated TCs. Summary IR digital data recorded from WSTF's APS testing is included within this document along with findings, lessons learned, and recommendations for further usage as a monitoring tool for the development of rocket engines.

Holleman, Elizabeth↗

NASA NTRS1966

Thermal Control of Liquid H2/O2 Rocket Engines and Vehicle Impact

Cryogenic propulsion systems are subjected to various engine start environments during ground testing, first burn, and restart after orbital coast. These operational modes necessitate careful engine-stage integration, particularly in thermal control of propellant feed systems and engine subsystems. The wide range of operational conditions imposed on propellant feed systems include varying heat loads and acceleration forces for static test first burn and orbital restart. One method used to provide Net Positive Suctign,Head (NPSH) and maintain turbopump I is thermal control propellant circulation through the feed system prior to engine start. The attendant design and operational criteria are inter-related with stage and vehicle systems, including for example, auxiliary propulsion and propellant ullaging methods such as propellant tank propulsive venting. Thermally sensitive engine components require environmental control, protection, or special sequencing in order to assure repeatable engine start performance. An example is thrust chamber pre-chill to avoid pump stall during rapid start transient. Trends in engine development are toward expanding system flexibility to accommodate the wide variety of environments without special pre-start conditioning.

Winstead, Thomas W.↗

NASA NTRS1966

Propulsion system selection considerations for the Voyager spacecraft.

Propulsion system selection considerations for Voyager spacecraft including Minuteman solid with liquid restartable midcourse, orbit trim system, LEM descent stage and transtage

SPACECRAFT PROPULSION↗

NASA NTRS2008

The Stationary Condensation and Radial Outflow of a Liquid Film on a Horizontal Disk

The application of capillary screen liquid acquisition devices to space-based cryogenic propulsion systems is expected to necessitate thermodynamic conditioning in order to stabilize surface tension retention characteristics. The present results have been obtained in the framework of the research of low gravity condensation-flow processes for conditioning cryogenic liquid acquisition devices. The following system is studied: On the top of a subcooled horizontal disk, a liquid film condenses from the ambient saturated vapor. The liquid is forcedly removed at the disk edge, and there is an outward radial flow of the film. Stationary regimes of the flow are uncovered such that (i) the gravity is negligible, being eclipsed by the capillary forces; (ii) the film thickness is everywhere much smaller than the disk radius; and (iii) the slow-flow lubrication approximation is valid. A nonlinear differential equation for the film thickness as a function of the radial coordinate is obtained. The (two-dimensional) fields of velocities, temperature and pressure in the film are explicitly determined by the radial profile of its thickness. The equilibrium is controlled by two parameters: (i) the vapor-disk difference of temperatures and (ii) the liquid exhaust rate. For the flow regimes with a nearly uniform film thickness, the governing equation linearizes, and the film interface is analytically predicted to have a concave-up quartic parabola profile. Thus, perhaps counter-intuitively, the liquid film is thicker at the edge and thinner at the center of the disk.

Bolshinskiy, Leonid↗

NASA NTRS2003

Microfabricated Liquid Rocket Motors

Under NASA Glenn Research Center sponsorship, MIT has developed the concept of micromachined, bipropellant, liquid rocket engines. This is potentially a breakthrough technology changing the cost-performance tradeoffs for small propulsion systems, enabling new applications, and redefining the meaning of the term low-cost-access-to-space. With this NASA support, a liquid-cooled, gaseous propellant version of the thrust chamber and nozzle was designed, built, and tested as a first step. DARPA is currently funding MIT to demonstrate turbopumps and controls. The work performed herein was the second year of a proposed three-year effort to develop the technology and demonstrate very high power density, regeneratively cooled, liquid bipropellant rocket engine thrust chamber and nozzles. When combined with the DARPA turbopumps and controls, this work would enable the design and demonstration of a complete rocket propulsion system. The original MIT-NASA concept used liquid oxygen-ethanol propellants. The military applications important to DARPA imply that storable liquid propellants are needed. Thus, MIT examined various storable propellant combinations including N2O4 and hydrazine, and H2O2 and various hydrocarbons. The latter are preferred since they do not have the toxicity of N2O4 and hydrazine. In reflection of the newfound interest in H2O2, it is once again in production and available commercially. A critical issue for the microrocket engine concept is cooling of the walls in a regenerative design. This is even more important at microscale than for large engines due to cube-square scaling considerations. Furthermore, the coolant behavior of rocket propellants has not been characterized at microscale. Therefore, MIT designed and constructed an apparatus expressly for this purpose. The report details measurements of two candidate microrocket fuels, JP-7 and JP-10.

Epstein, Alan H.↗

NASA NTRS2006

Nondestructive Evaluation of the Gimbal Joint Flowliner Slots in the Space Shuttle Main Propulsion System Hydrogen Feedline

Fatigue cracks were discovered in the STS-112 Liquid Hydrogen Feedline flowliners in 2002. This led to a development program aimed at providing nondestructive evaluation methods and techniques to verify the existence of these types of cracks in oval shaped slots cut into the ends of the feedlines above the bellows joints. These slots were used to improve flow dynamics and to facilitate cleaning in the bellow joint region. These types of fatigue cracks posed a possible metal debris ingestion threat for the Space Shuttle Main Engines, which attached to these particular joints. Results of this program produced three reliable inspection techniques utilizing the imaging of replisets with a Scanning Electron microscope, eddy current, and ultrasound. The program developed unique probes and fixtures and in the case of eddy current and ultrasound, provided qualification and certification of the particular techniques by various Design of Experiments and Probability of Detection studies utilizing multiple inspectors.

Suits, Michael W.↗

NASA NTRS2001

Nuclear Thermal Rocket (NTR) Propulsion and Power Systems for Outer Planetary Exploration Missions

The high specific impulse (I (sub sp)) and engine thrust generated using liquid hydrogen (LH2)-cooled Nuclear Thermal Rocket (NTR) propulsion makes them attractive for upper stage applications for difficult robotic science missions to the outer planets. Besides high (I (sub sp)) and thrust, NTR engines can also be designed for "bimodal" operation allowing substantial amounts of electrical power (10's of kWe ) to be generated for onboard spacecraft systems and high data rate communications with Earth during the course of the mission. Two possible options for using the NTR are examined here. A high performance injection stage utilizing a single 15 klbf thrust engine can inject large payloads to the outer planets using a 20 t-class launch vehicle when operated in an "expendable mode". A smaller bimodal NTR stage generating approx. 1 klbf of thrust and 20 to 40 kWe for electric propulsion can deliver approx. 100 kg using lower cost launch vehicles. Additional information is contained in the original extended abstract.

Borowski, S. K.↗

NASA NTRS1975

Pressurization systems for liquid rockets

Guidelines for the successful design of pressurization systems for main propulsion, auxiliary propulsion, and attitude control systems for boosters, upper stages, and spacecraft were presented, drawing on the wealth of design experience that has accumulated in the development of pressurization systems for liquid rockets operational in the last 15 years. The design begins with a preliminary phase in which the system requirements are received and evaluated. Next comes a detail-design and integration phase in which the controls and the hardware components that make up the system are determined. The final phase, design evaluation, provides analysis of problems that may arise at any point in the design when components are combined and considered for operation as a system. Throughout the monograph, the design tasks are considered in the order and manner in which the designer must handle them.

Source record↗

NASA NTRS1989

Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Appendix B: Liquid rocket booster acoustic and thermal environments

The ascent thermal environment and propulsion acoustic sources for the Martin-Marietta Corporation designed Liquid Rocket Boosters (LRB) to be used with the Space Shuttle Orbiter and External Tank are described. Two designs were proposed: one using a pump-fed propulsion system and the other using a pressure-fed propulsion system. Both designs use LOX/RP-1 propellants, but differences in performance of the two propulsion systems produce significant differences in the proposed stage geometries, exhaust plumes, and resulting environments. The general characteristics of the two designs which are significant for environmental predictions are described. The methods of analysis and predictions for environments in acoustics, aerodynamic heating, and base heating (from exhaust plume effects) are also described. The acoustic section will compare the proposed exhaust plumes with the current SRB from the standpoint of acoustics and ignition overpressure. The sections on thermal environments will provide details of the LRB heating rates and indications of possible changes in the Orbiter and ET environments as a result of the change from SRBs to LRBs.

Source record↗